1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a stator vane with rim cavity seal.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, a compressor provides compressed air into a combustor in which a fuel is burned to produce a hot gas flow. The hot gas flow is passed through a turbine to convert the heat energy from the hot gas flow into mechanical energy that is used to power the compressor and, in the case of an industrial gas turbine (IGT) engine, to drive an electric generator. In a large IGT, efficiency is major priority in order to provide the highest electrical output to fuel cost ratio possible. The turbine includes a number of stages of stator vanes and rotor blades in which rotary seals are used between parts to prevent the hot gas flow from leaking around blade tips or from passing into areas sensitive to high temperatures.
One problem with today's IGT engines is the ability to make improvements to an engine that is difficult to make design changes on. The stator vanes in the turbine section require a seal between the inner shroud portion and the two rotor blades on either sides of the vane. U.S. Pat. No. 6,761,526 B2 issued to Soechting et al on Jul. 13, 2004 and entitled COOLING STRUCTURE OF STATIONARY BLADE, AND GAS TURBINE show (in FIG. 1 of the Seochting patent) a seal formed on the inner end of the vane extending from a seal supporting part that forms the seal with two sealing arms that extend from the rotor disks of the blades on both sides of the vane. Because of thermal growths during engine transients (engine operation during startups and shut-downs) and steady state operations, the seal gap can vary considerably and produce a large opening for leakage across the seal. In this particular situation, the hot gas flow on the left or upstream side of the vane is at a higher pressure and higher temperature than on the downstream or right side of the vane. In order to prevent ingestion of the hot gas flow from the upstream side into the box rim cavity, more cooling air from the vane is required to be pumped into the cavity and is therefore wasted.
In the prior art, passive tip clearance control has been used in aero engines for the reduction of tip leakage control. Cooling air has been used in the cooling of the blade outer seal carrier to minimize the radial thermal expansion. This minimizes the radial tip clearance between the blade and the outer air seal. In addition, high effective cooling schemes were also incorporated into the turbine tip cooling and sealing designs for the reduction of leakage flow across the blade tip. In one prior art engine, the rotor shaft is moved axially by a hydraulic actuator in order to control the rotor blade tip clearance. However, very little progress has been made in the control of rim cavity leakage flow distribution for the reduction of the total purge air demand, especially for a large IGT design application. Due to the large pressure differential between the front rim cavities versus the aft rim cavity, the front rim cavity requires a higher purge air pressure than the aft rim cavity to prevent the hot gas ingestion into the forward cavity. Cooling air for both the forward and the aft rim cavities is provided from the same source, the inter-stage seal housing. An open gap in-between the seal housing versus the rotor will result in purge air being distributed unevenly. A majority of the purge air is passed through the sealing gap and exits from the aft rim cavity. In some cases, hot gas ingestion into the front rim cavity will result from the purge air uneven distribution.
In some IGT engines, the rotor disk cannot withstand exposure to a temperature above 450C because of the thermal properties of the shaft. Higher prolonged temperature exposure due to hot gas flow leakage will result in decreased life of the part from crack growth. Excess cooling air flow to the box rim cavity is required to prevent over-temperature of the shaft. Thus, there is a need in the prior art to improve on the seal capability within the turbine to prevent exposure of certain parts from thermal exposure in order to prolong the useful life of these parts.